Active measurement of gas flow velocity or simultaneous measurement of velocity and temperature, including in gas turbine combustors

ABSTRACT

Active acoustic velocity and pyrometry-based gas flow velocity and temperature measurement, such as for monitoring of gas turbine combustors, including industrial gas turbine (IGT) combustors is incorporated into the combustion monitoring and control system by addition of an acoustic transmitter or acoustic transceiver that transmits a sound wave in a line-of-sight with a plurality of acoustic sensors, such as dynamic pressure sensors. For velocity measurement, sound transmission time-of-flight that is directed generally along the gas flow path is measured by the controller and correlated with gas flow velocity along the line-of-sight. Similarly, sound transmission time-of-flight is correlated with temperature along the line-of-sight. Path(s) of acoustic transmission serve as velocity or velocity/absolute temperature measurement. In an integrated thermoacoustic pressure-based sensor and monitoring/control system embodiment, the controller correlates velocity and, if desired, absolute active path temperatures with acoustic transmission and time-of-flight analysis techniques.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation in part of copending United Statespatent application entitled “Active Temperature Monitoring In GasTurbine Combustors”, filed on Dec. 18, 2013, Ser. No. 14/132,001.

This application incorporates by reference the following co-pendingUnited States utility patent applications in their entirety as if fullyset forth herein:

“Active Measurement Of Gas Flow Temperature, Including In Gas TurbineCombustors”, filed concurrently herewith, Serial Number unknown, DocketNumber 2014P05241US;

“Multi-Functional Sensor System For Gas Turbine Combustion MonitoringAnd Control” filed on Dec. 18, 2013, Ser. No. 14/109,992;

“Temperature Measurement In A Gas Turbine Engine Combustor”, filed onMar. 14, 2013, Ser. No. 13/804,132; and

“Gas Turbine Engine Control Using Acoustic Pyrometry”, filed on Dec. 14,2010, Ser. No. 12/967,148, Publication No. US2012/0150413.

This application also incorporates by reference in its entirety as iffully set forth herein U.S. Pat. No. 7,853,433, “Combustion AnomalyDetection Via Wavelet Analysis Of Dynamic Sensor Signals”, issued Dec.14, 2010.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The invention relates to active measurement of gas flow velocity orsimultaneous measurement of velocity and temperature, such as combustiongas flow in combustors of gas turbine engines. Such engines include, byway of example, industrial gas turbine (IGT) engines, other types ofstationary gas turbine, marine, aero and other vehicular gas turbineengines. More particularly, embodiments of velocity or simultaneousvelocity/temperature measurement methods and apparatus disclosed hereinutilize a common sensing and control system for combustor velocity andtemperature determination. In embodiments disclosed herein acircumferential and axially spaced array of acoustic sensors, acoustictransmitters and/or transceivers are utilized for one or more ofreal-time active combustor gas flow velocity or simultaneousvelocity/temperature measurement. Gas flow velocity and temperature dataare used for engine combustion monitoring and control.

2. Description of the Prior Art

Combustion turbines, such as gas turbine engines for any end useapplication, generally comprise a compressor section, a combustorsection, a turbine section and an exhaust section. In operation, thecompressor section inducts and compresses ambient air. The combustorsection generally may include a plurality of combustors for receivingthe compressed air and mixing it with fuel to form a fuel/air mixture.The fuel/air mixture is combusted by each of the combustors to form ahot working gas that may be routed to the turbine section where it isexpanded through alternating rows of stationary airfoils and rotatingairfoils and used to generate power that can drive a rotor. Theexpanding gas exiting the turbine section can be exhausted from theengine via the exhaust section.

Combustion anomalies, such as flame flashback, have been known to occurin combustion sections of gas turbine engines. Flame flashback is alocalized phenomenon that may be caused when a turbulent burningvelocity of the air and fuel mixture exceeds an axial flow velocity inthe combustor assembly, thus causing a flame to anchor onto one or morecomponents in/around the combustor assembly, such as a liner disposedaround the combustion chamber. The anchored flame may burn through thecomponents if a flashback condition remains for extended periods of timewithout correction thereof. Thus, flame flashback and/or othercombustion anomalies may cause undesirable damage and possibly evendestruction of combustion engine components, such that repair orreplacement of such components may become necessary.

The fuel/air mixture at the individual combustors is controlled duringoperation of the engine to maintain one or more operatingcharacteristics within a predetermined range, such as, for example, tomaintain a desired efficiency and/or power output, control pollutantlevels, prevent pressure oscillations and prevent flameouts. In a knowntype of control arrangement, a bulk turbine exhaust temperature may alsobe monitored as a parameter that may be used to monitor the operatingcondition of the engine. For example, a controller may monitor ameasured turbine exhaust temperature, and a measured change intemperature at the exhaust may result in the controller changing anoperating condition of the engine. In other known types of controlarrangements discrete pitot-static or multi hole pressure probes areutilized to determine gas flow velocity at specific locations, but gridarrays of such probes disrupt gas flow and introduce measurement errors.Due to such gas flow disruptions grid arrays, when employed, havelimited numbers of widely spaced probes, which provide relatively coarsegas flow velocity distribution and profile information.

At present, there are several different types of sensors and sensingsystems that are being used in the industry for monitoring combustionand maintaining stability of the combustion process for engineprotection. For example, dynamic pressure sensors are being used forcombustion stability and resonance control. Passive visual (opticalvisible light and/or infrared spectrum) sensors, ion sensors and GeigerMueller detectors are used to detect flame on/off in the combustor,while thermocouples are being used for flashback detection. With respectto known combustion gas flow velocity (u) monitoring methods,pitot-static and multi hole pressure probes utilize differentialpressure techniques, hot wire probes utilize thermal anemometrytechniques, while Laser Doppler and Particle Image Velocimetry systemsutilize optical techniques to characterize gas flow velocities.Differential pressure and thermal anemometry instruments are intrusivepoint measurement devices that disturb local gas flow around theinstruments. Laser Doppler and Particle Image Velocimetry instrumentsrespectively provide non-intrusive point and 2- or 3-dimensionalnon-intrusive gas flow velocity measurement although they both requireparticle seeding of the flow. In addition, sophisticated laser basedmeasurements such as Filtered Rayleigh Scattering (FRS) and other suchlaser spectroscopy based techniques have been deployed to measure gasvelocity. However, these techniques are more complex than intrusivedifferential pressure or thermal anemometry instruments and require morespecialized training to implement in monitoring systems. Moreover, mostoptical techniques for velocity are geared towards laboratoryenvironments rather than in operative engines at power plant fieldsites. With respect to temperature (T) monitoring techniques, knownRaman Spectroscopy, Laser Induced Fluorescence (for both u and Tmonitoring), and Coherent Anti-Stokes Raman Spectroscopy (CARS) (forboth u and T monitoring) instrumentation systems are also intended forlaboratory environments, rather than for field use in fossil powergeneration equipment. Tunable Diode Laser Absorption Spectroscopy(TDLAS) instrumentation is used in some industrial power generationfield applications, such as for temperature measurement in boilers butthat instrumentation is extremely costly: approximately US $500,000 persystem. Other types of temperature measurement and combustion anomalydetection systems have had greater acceptance in power generationindustry field applications.

Particularly, U.S. Pat. No. 7,853,433 detects and classifies combustionanomalies by sampling and subsequent wavelet analysis of combustorthermoacoustic oscillations representative of combustion conditions withsensors, such as dynamic pressure sensors, accelerometers, hightemperature microphones, optical sensors and/or ionic sensors. UnitedStates Publication No. US2012/0150413 utilizes acoustic pyrometry in anIGT exhaust system to determine upstream bulk temperature within one ormore of the engine's combustors. Acoustic signals are transmitted fromacoustic transmitters and are received by a plurality of acousticreceivers. Each acoustic signal defines a distinct line-of-sound pathbetween a corresponding transmitter and receiver pair. Transmittedsignal time-of-flight is determined and processed to determine a pathtemperature. Multiple path temperatures can be combined and processed todetermine bulk temperature at the measurement site. The determined pathor bulk temperature or both can be utilized to correlate upstreamtemperature in the combustor. Co-pending United States utility patentapplication Ser. No. 13/804,132 calculates bulk temperature within acombustor, using a so-called dominant mode approach, by identifying anacoustic frequency at a first location in the engine upstream from theturbine (such as in the combustor) and using the frequency fordetermining a first bulk temperature value that is directly proportionalto the acoustic frequency and a calculated constant value. A calibrationsecond temperature of the working gas is determined in a second locationin the engine, such as the engine exhaust. A back calculation isperformed with the calibration second temperature to determine atemperature value for the working gas at the first location. The firsttemperature value is compared to the back calculated temperature valueto change the calculated constant value to a recalculated constantvalue. Subsequent first temperature values at the combustor may bedetermined based on the recalculated constant value.

Thus, different adverse conditions related to combustion gas flowtemperature, anomalies and velocity currently require separate sensordesigns and/or separate sensing systems to detect those conditions.Known combined IGT and other types of gas turbine engine monitoring andcontrol system sensor and detection approaches have not covered allpossible adverse combustion fault detections. Installation of differenttypes of disparate sensors and sensing systems in a single combustionturbine engine increases installation cost and maintenance expense.Also, the disparate sensors and sensing systems inherently introduceresponse lags and delays in the overall engine control system.

Thus, a need exists in the art for an integrated gas turbine enginemonitoring and control system for measuring gas flow velocity,temperature and detecting a broad range of possible combustor failuresor, more satisfactorily precursors to faults, during combustion, sharingcommon sensors and, if desired, a common controller.

Another need exists in the art for a gas turbine engine active velocityand temperature monitoring system that determines actual combustortemperature in real time without the need to obtain referencetemperatures from other locations within the engine, such as known bulktemperature systems that back calculate combustor temperature based ontemperature measurements obtained in the engine exhaust system.

An additional need exists for an active gas flow velocity andtemperature monitoring system that shares sensors commonly used withcombustion turbine monitoring and control systems, so that activevelocity and temperature monitoring can be integrated within themonitoring and control system.

SUMMARY OF THE INVENTION

An object of embodiments of the invention is creation of an active gasflow velocity monitoring system that shares sensors commonly used withgas turbine monitoring and control systems, so that active gas flowvelocity monitoring can be integrated within the monitoring and controlsystem.

An additional object of embodiments of the invention is creation of anactive simultaneous gas flow velocity and temperature monitoring systemthat shares sensors commonly used with gas turbine monitoring andcontrol systems, so that active simultaneous gas flow velocity andtemperature monitoring can be integrated within the monitoring andcontrol system.

Another object of embodiments of the invention is creation of anintegrated gas turbine monitoring and control system for detecting abroad range of possible combustor failures or, more satisfactorilyprecursors to faults, during combustion, sharing common sensors and, ifdesired, a common controller with the velocity or velocity/temperaturemeasurement system.

Another object of embodiments of the invention is creation of a gasturbine active temperature monitoring system that determines actualcombustor temperature in one or more combustors in real time without theneed to obtain reference temperatures from other locations within theengine, such as known bulk temperature systems that back calculatecombustor temperature based on temperature measurements obtained in theengine exhaust system.

These and other objects are achieved in one or more embodiments of theinvention by active acoustic velocity and pyrometry-based gas flowvelocity and temperature measurement systems and methods describedherein. Embodiments of the velocity or velocity/temperature monitoringsystems and methods are used for monitoring combustion gas within gasturbine combustors, including industrial gas turbine (IGT) combustors.These velocity or velocity/temperature monitoring systems and methodsare is incorporated into the turbine combustion monitoring and controlsystem by addition of at least one acoustic transmitter or acoustictransceiver that transmits a sound wave in a line-of-sight with at leastone or a plurality of acoustic sensors, such as dynamic pressuresensors. For velocity measurement, sound transmission time-of-flightthat is directed generally transverse to the gas flow path is measuredby the controller and correlated with gas flow velocity along theline-of-sight. Similarly, sound transmission time-of-flight iscorrelated with temperature along the line-of-sight. Path(s) of acoustictransmission serve as velocity or velocity/absolute temperaturemeasurement. In an integrated acoustic pressure-based sensor andmonitoring/control system embodiment, the controller correlates velocityand, if desired, absolute active path temperatures with acoustictransmission and time-of-flight analysis techniques. In this way acombustion monitoring and control system that incorporates embodimentsof the invention can identify and classify combustion anomalies andactively control the gas turbine combustion process within the enginecombustors.

Optionally acoustic pyrometry-based active temperature monitoring isincorporated into the monitoring and control system, alone or with otherembodiments of the invention described herein, by addition of anacoustic transmitter or acoustic transceiver that transmits a sound wavein a line-of-sight with at least one paired acoustic sensor or aplurality of acoustic sensors. Sound transmission time-of-flight ismeasured by the controller and correlated with path temperature alongthe line-of-sight. The active path temperatures may be processed by thecontroller to determine combustor bulk temperature. The path(s) ofacoustic transmission across the combustor gas path serve as theabsolute temperature measurement. In some embodiments the combustoracoustic pyrometry system is used for calibrating the passive bulktemperature measurement using the dominant mode approach. Thus theactive path temperature determined by acoustic pyrometry methods can beutilized as the second or reference temperature reading for calibratingand corroborating acoustic frequency based bulk temperature monitoringsystems, when they are incorporated in a common control and monitoringsystem.

In some embodiments all of the monitoring functions of the integratedmonitoring and control system can be performed with a commonly sharedarray of acoustic sensors that function as both combustion dynamicsthermoacoustic vibration/wave receivers and acoustic transmitters. Insuch an integrated thermoacoustic pressure-based sensor andmonitoring/control system, the controller correlates performance ofcombustion thermoacoustic properties in order to identify combustionanomalies by wavelet or Fourier analysis techniques; determine bulktemperature characteristics within the combustor with dominant modefrequency analysis techniques; or determine gas flow relative orabsolute velocity, as well as active path absolute temperatures, withinthe combustor with acoustic transmission and time-of-flight analysistechniques described herein.

Embodiments of this monitoring and control system with multi-functionsensors operate by evaluating dynamic pressure sensor data forcombustion anomalies while also monitoring accurate combustion gasvelocity or velocity/temperature continuously in real time. Once themonitoring and control system detects that there is a fault anomaly itevaluates the fault in conjunction with the monitored combustor gasvelocity or velocity/temperature to classify or determine the type offault. This system classifies various types of developing combustionanomalies, with the ability to predict flame on, flame out, flashback.The system embodiments also monitor vibratory responses and resonances(low, intermediate and high frequency dynamics) of the combustor basket.As mentioned, accurate real time, continuous active gas flow velocity orvelocity/absolute path temperature monitoring and/or bulk meantemperature monitoring by embodiments of the monitoring/control systemare available for optimizing engine control and performance. Embodimentsof the system of the present invention are implemented in any type ofknown combustor structure, whether of can, can-annular or annularconstruction, to control and regulate the gas turbine combustion controlprocess during engine operation.

Embodiments of the invention are also directed to methods for activelymonitoring gas flow velocity by placing at least one first acoustictransmitter and at least one first acoustic sensor respectively orientedin upstream and downstream transverse locations in a gas flow path andin a distinct first line-of-sound path relative to each other, with thefirst sensor capable of generating first sensor output signalsindicative of first thermoacoustic oscillations. The at least one firsttransmitter and the at least one first sensor are coupled to acontroller that is capable of causing the at least one first transmitterto transmit first acoustic signals within the gas flow path and capableof correlating first sensor output signals with gas flow velocity. Firstacoustic signals are transmitted from the at least one first acoustictransmitter. The first acoustic signals from the at least one firstacoustic transmitter and are received by the at least one first acousticsensor, which in turn generates first dynamic sensor output signals thatincludes contributions of the received first acoustic signals. Thecontroller determines a first time-of-flight for the first acousticsignals traveling along each of the first line of sound paths andprocesses the first time-of-flight for the first acoustic signalstraveling along their first line of sound paths to determine respectivegas flow velocity along each respective first line-of-sound path. Insome embodiments the processing to determine absolute gas flow velocityincludes compensation for temperature, gas constant and speed of soundthermodynamic influences on the first time-of-flight.

In some embodiments the temperature compensation thermodynamicinfluences on the first time-of-flight is performed by activelymonitoring temperatures in the gas flow by placing at least one secondacoustic transmitter and at least one second acoustic sensorrespectively oriented in a common axial plane in the gas flow path andin a distinct second line-of-sound path relative to each other andcoupling both to the controller. The second sensor is capable ofgenerating second sensor output signals indicative of secondthermoacoustic oscillations. The controller is capable of causing the atleast one second transmitter to transmit second acoustic signals withinthe gas flow path and capable of correlating the second sensor outputsignals time-of-flight with gas flow temperature. The controller thenprocesses the second time-of-flight for the second acoustic signalstraveling along the second line of sound paths to determine respectivegas flow temperature along each respective second line-of-sound path.

In other embodiments speed of sound compensation for thermodynamicinfluences on the first time-of-flight is performed rather thantemperature compensation. In these embodiments compensation for thespeed of sound in the velocity monitoring is accomplished bysubstituting for the first transmitters a set of first transducers thatare capable of transmitting and receiving acoustic signals, andgenerating output signals and substituting for the first sensors secondset of transducers that are capable of transmitting and receivingacoustic signals and generating output signals. The respective at leastone of the respective first and second transducers is coupled to thecontroller, which in turn is capable of causing either to transmit firstacoustic signals within the gas flow path and capable of correlatingtransducer output signals time-of-flight with gas flow velocity. In thisembodiment first acoustic signals are transmitted from the at least onefirst transceiver/transducer and received by the secondtransceiver/transducer, thereby generating first dynamic sensor outputsignals with the at least one second transceiver/transducer thatincludes contributions of the received first acoustic signals. Reversedfirst acoustic signals are transmitted from the at least one secondtransceiver/transducer and received by the first transceiver/transducer,which generates first reversed dynamic sensor output signals thatincludes contributions of the received reversed first acoustic signals.The controller is used to determine a first time-of-flight for therespective first and first reversed acoustic signals traveling alongeach of the first line of sound paths; and in turn processes the firsttime-of-flight for the respective first and reversed acoustic signalstraveling along their respective first line of sound paths to determinethe speed of sound c.

Other embodiments of the invention are directed to systems formonitoring gas flow velocity in a turbine engine, including at least onefirst acoustic transmitter and at least one first acoustic sensorrespectively oriented in upstream and downstream transverse locations ina gas flow path of a turbine engine and in a distinct firstline-of-sound path relative to each other. The first sensor is capableof generating first sensor output signals indicative of firstthermoacoustic oscillations. A controller is coupled to the at least onefirst transmitter and the at least one first sensor that is capable ofcausing the at least one first transmitter to transmit first acousticsignals within the gas flow path and capable of correlating first sensoroutput signals time-of-flight with gas flow velocity. The correlation isperformed with the controller by: transmitting first acoustic signalsfrom the at least one first acoustic transmitter and receiving the firstacoustic signals from the at least one first acoustic transmitter thatin turn generates first dynamic sensor output signals with the at leastone first acoustic sensor that includes contributions of the receivedfirst acoustic signals. The controller determines a first time-of-flightfor the first acoustic signals traveling along each of the first line ofsound paths and processes the first time-of-flight to determinerespective gas flow velocity along each respective first line-of-soundpath. In other embodiments the system also simultaneously monitorsactive gas flow temperature, with addition of at least one secondacoustic transmitter and at least one second acoustic sensorrespectively oriented in a common axial plane in the turbine gas flowpath and in a distinct second line-of-sound path relative to each other.The second sensor is capable of generating second sensor output signalsindicative of second thermoacoustic oscillations. The controller iscoupled to both the second transmitter and second sensor: it is capableof causing the at least one second transmitter to transmit secondacoustic signals within the gas flow path and processing the secondtime-of-flight for the second acoustic signals traveling along thesecond line of sound paths to determine respective gas flow temperaturealong each respective second line-of-sound path. The determinedrespective gas flow temperatures are used to compensate for temperatureeffects on the first time-of-flight. Alternatively speed of soundcompensation for thermodynamic influences on the gas flow velocity areutilized to determine absolute gas flow velocity, as discussed above.

Additional embodiments of the invention are directed to a gas turbineengine apparatus, having compressor section; a combustor sectionincluding a plurality of combustors, each combustor having an injectorsystem for regulating fuel/air mixture a turbine section; and a systemfor monitoring gas flow velocity in the turbine engine. The systemincludes at least one first acoustic transmitter and at least one firstacoustic sensor respectively oriented in upstream and downstreamtransverse locations in a gas flow path of a turbine engine and in adistinct first line-of-sound path relative to each other. The firstsensor is capable of generating first sensor output signals indicativeof first thermoacoustic oscillations. A controller is coupled to the atleast one first transmitter and the at least one first sensor that iscapable of causing the at least one first transmitter to transmit firstacoustic signals within the gas flow path and that is capable ofcorrelating first sensor output signals with gas flow velocity. The gasflow velocity is correlated by transmitting first acoustic signals fromthe at least one first acoustic transmitter and receiving the firstacoustic signals from the at least one first acoustic transmitter andgenerating first dynamic sensor output signals with the at least onefirst acoustic sensor that includes contributions of the received firstacoustic signals. A first time-of-flight for the first acoustic signalstraveling along each of the first line of sound paths is determined. Thefirst time-of-flight for the first acoustic signals traveling alongtheir first line of sound paths is processed to determine respective gasflow velocity along each respective first line-of-sound path. Asdescribed above, the processing to determine respective gas flowvelocity in some embodiments includes compensation for temperature, gasconstant and speed of sound thermodynamic influences on the firsttime-of-flight.

The respective objects and features of the present invention may beapplied jointly or severally in any combination or sub-combination bythose skilled in the art.

BRIEF DESCRIPTION OF THE DRAWINGS

The teachings of the present invention can be readily understood byconsidering the following detailed description in conjunction with theaccompanying drawings, in which:

FIG. 1 is a perspective cross-sectional view of a gas turbine engineillustrating implementation of a system for determining combustor gasflow active velocity and temperature measurement, in accordance withembodiments of the invention;

FIG. 2 is a cross-sectional view of a gas turbine combustorincorporating an embodiment of a monitoring system for determiningcombustor gas flow active velocity and temperature measurement, inaccordance with embodiments of the invention;

FIG. 3 is a cross-sectional view of the system of FIG. 2, taken along3-3 thereof, in accordance with aspects of the invention;

FIG. 4 is a block diagram of an embodiment of a controller forimplementing embodiments of the present invention in the monitoringsystem for determining combustor gas flow active velocity andtemperature measurement, in accordance with embodiments of theinvention;

FIG. 5 is a schematic perspective view of exemplary sonic sensor arraysused by the gas flow monitoring system to measure gas flow velocity in agas turbine combustor, in accordance with embodiments of the invention;

FIG. 6 is an exemplary schematic representation of gas flow velocity inthe turbine combustor of FIG. 5 in the line-of-sight between acousticsensors 32B and 34C;

FIG. 7. is a cross-sectional slice A of the gas flow velocity of FIG. 6taken along 7-7 thereof, which corresponds to the line-of-sight betweenacoustic sensors 32B and 34C;

FIG. 8 is a composite gas flow velocity profile of the respectivevelocities measured by the gas flow velocity monitoring system, inaccordance with embodiments of the invention;

FIG. 9 is a schematic perspective view of exemplary sonic sensor arraysused to measure gas flow temperature in a gas turbine combustor, inaccordance with embodiments of the invention;

FIG. 10 is a flow chart illustrating implementation of an embodiment ofthe methods for measuring gas flow velocity and temperature activemeasurement in a gas turbine combustor, in accordance with embodimentsof the invention; and

FIG. 11 is a flow chart illustrating implementation of an embodiment ofthe method for measuring active gas flow velocity, in accordance withembodiments of the invention.

To facilitate understanding, identical reference numerals have beenused, where possible, to designate identical elements that are common tothe figures.

DETAILED DESCRIPTION

After considering the following description, those skilled in the artwill clearly realize that the teachings of the invention can be readilyutilized for active acoustic velocity and pyrometry-based gas flowvelocity and temperature measurement. Embodiments of the invention areused for monitoring of gas turbine combustors, including industrial gasturbine (IGT) combustors by incorporating them into the combustionmonitoring and control system by addition of an acoustic transmitter oracoustic transceiver that transmits sound waves through gas flow in aline-of-sight with a plurality of acoustic sensors, such as dynamicpressure sensors. For velocity measurement, sound transmissiontime-of-flight that is directed generally transversely through the gasflow path is measured by the controller and correlated with gas flowvelocity along the line-of-sight. The gas flow velocity determinationincludes compensation for impact of the thermodynamically interrelatedtemperature, gas constant and speed of sound influences on the firsttime-of-flight, in order to determine absolute gas flow velocity.

In an integrated acoustic pressure-based sensor and monitoring/controlsystem embodiment, the controller correlates velocity and, if desired,absolute active path temperatures simultaneously with acoustictransmission and time-of-flight analysis techniques. Where velocity andtemperature are measured simultaneously the absolute active pathtemperature is utilized to compensate for the aforementionedthermodynamic influences on gas flow absolute velocity. Alternatively inother embodiments the speed of sound influence on the firsttime-of-flight is utilized to determine absolute gas flow velocityrather than absolute active path temperature. In such embodiments,compensation for the speed of sound in the velocity monitoring isaccomplished by substituting for the first transmitters a set of firsttransceiver/transducers that are capable of transmitting and receivingacoustic signals, and generating output signals and substituting for thefirst sensors a set of second transducers that are capable oftransmitting and receiving acoustic signals and generating outputsignals. Acoustic signals are transmitted and received from the first tothe second transducers and time-of-flight is determined. A reverseacoustic signal is transmitted from the second to the first transducersand the reverse time-of-flight is determined. The respective first andfirst reversed acoustic signals times-of-flight are used to determinethe speed of sound c. The determined speed of sound c is then utilizedfor determination of the actual gas flow velocity.

In embodiments of the invention active velocity or activevelocity/temperature measurements are used as monitoring parameters forgas flow in a combustion monitoring and control system that can identifyand classify gas flow anomalies (e.g., combustion anomalies), forexample by using wavelet or Fourier analysis techniques. Someembodiments of the methods and system incorporate one or more acousticdynamic pressure transceiver/transducer combination transmitter/sensorsthat are selectively oriented or arrayed in sequential axial planarpositions within the combustor. Known transceiver/transducer componentdesigns and their related controller components have been used reliablyand cost effectively in the past in power generation field service. Byreconfiguring those types of known components into the gas flow controland monitoring systems of the present invention combustion turbine andother combustion power generation equipment can be monitored andcontrolled with simpler instrumentation hardware configurations thatprovide detailed active gas flow velocity and temperature distributioninformation useful for precise combustion control.

Monitoring and Control System Structure

Referring to FIGS. 1 and 2 an exemplary industrial gas turbine engine 10is shown. The exemplary engine 10 includes a compressor section 12, acombustor section 14, a turbine section 16, and an exhaust section orsystem 18. The combustor section 14 includes a plurality of combustors20. Each combustor 20 has a combustion shell 22 and a cover plate 24.The combustor liner or basket 26 and transition duct 27 define a passagefor conveying hot working gas that flows in the direction F to theturbine section 16. The system of the present invention is operable withknown combustor geometry gas turbine engine designs, including can,can-annular or annular construction combustors in stationary land-basedor vehicular applications.

During operation of the engine 10, compressed air from the compressorsection 12 is provided to the combustor section 14 where it is combinedwith fuel supplied by fuel injection system 28 in the combustors 14. Thefuel/air mixture is ignited to form combustion products comprising thehot working gas. It may be understood that combustion of the fuel andair may occur at various axial locations along the passage through thecombustor liner or basket 26 and the transition duct 27 to the inlet ofthe turbine section 16. The hot working gas is expanded through theturbine section 16 and is exhausted through the exhaust section/system18.

Referring to FIGS. 1 and 2, in accordance with an aspect of theinvention, a combustion monitoring and control system 29 is provided,which can identify and classify combustion anomalies and activelycontrol the gas turbine combustion process within one or more of theengine 10 combustors 20. In this regard, the engine 10 may include maycomprise one or more of the monitoring and control system(s) 29: e.g.,one system 29 for each combustor 20, or a single system 29 may serviceeach combustor 14 of the engine 10. Similarly, clusters of combustors 20may be served by one system 29, with other cluster(s) being served byother systems. Thus the consolidated monitoring system for an engine 10can determine deviations between respective combustors and compare theirrelative performance no matter what engine combustor structure ororientation is employed by the engine design: whether a stationary,land-based turbine engine or a vehicular engine for aero, marine or landvehicular applications.

As shown in FIGS. 2, 3, 5 and 9 the system 29 includes an array of aplurality of known acoustic transceiver/transducers 32A-H and 34A-H thatare capable of transmitting and receiving acoustic oscillation wavesalong exemplary the line-of-sight paths shown in dashed lines in FIGS. 5and 9. The transceiver/transducer arrays 32, 34 are capable ofgenerating respective sensor output signals indicative of combustionthermoacoustic oscillations in each respective monitored and controlledcombustor 20. Other system embodiments can be constructed with at leasttwo, but preferably more acoustic sensors, whether functionally part ofa transceiver component or as a stand-alone component. Acousticfrequencies and amplitudes sensed by those acoustic sensor portions ofthe transceivers are generated as a result of combustion events in theworking combustion gas, defining acoustic sources that occur within thecombustor 20 hot gas path. The monitoring and control system 29 isconfigured to transform the sensed thermoacoustic oscillationinformation into a form that enables the occurrence of combustionanomalies of interest to be discerned. As such, flame flashback eventsand other types of combustion anomalies of interest may be detected andextracted from sensed thermoacoustic oscillations in the combustor 14that are monitored by the transceiver/transducer/sensors positioned inand/or around the combustor 14. Depending upon the system 29configurations and application, the acoustic sensors comprise anycombination of one or more of a dynamic pressure sensor, a microphone,an optical sensor or an ionic turbine inlet sensor. Pressure sensorssense the amplitudes of thermoacoustic oscillations in the combustor 20as well as pulsation frequencies. A high temperature microphone may beutilized to measure acoustic fluctuations in the combustor 14. Anoptical sensor may be utilized to measure a dynamic optical signalwithin the combustor 20. An ionic sensor may be utilized to measuredynamic ionic activity within the combustor 20.

An exemplary acoustic sensor array shown schematically in FIGS. 2, 3, 5and 9 comprises transceiver/transducers 32A-H and 34A-H that function asat least one acoustic transmitter that transmits in turn to at least oneand preferably a plurality of the dynamic pressure sensors in the array.The transceiver/transducers 32, 34 are arrayed axially and radiallywithin the combustor 20 by known mounting structures and methods, suchas J tubes or rakes, within the combustor shell 22 proximal thecombustor basket or liner 26, and/or proximal the transition 27 junctionwith the turbine section 16. In FIG. 3 the sensors areradially/circumferentially arrayed transceivers 34A-34F1 that arecapable of transmitting and receiving acoustic oscillation waves alongthe line-of-sight paths similar to the transceivers 32A-H shown indashed lines in FIG. 9. Other types of known sensors, such as individualthermocouple temperature sensors or thermocouple arrays may be employedwithin the gas turbine engine. For example in FIG. 3 thermocouple 36measures combustion temperature in the combustor 20. While exemplarythree-dimensional annular combustion flow paths and axially spaced,two-dimensional circular-annular transceiver/transducer arrays are shownin the figures, other combustion flow path and array orientations may beutilized, in practicing embodiments of the invention, including square-or rectangular-shaped geometries.

As shown in greater detail in FIGS. 3 and 4, the monitoring and controlsystem 29 comprises a known controller 40, coupled to thetransceiver/transducers 32, 34, that is capable of correlating sensoroutput signals with gas flow velocity and combustion temperature in amonitoring section 42 and conducting combustion dynamics analysis of thecombustion process in an analysis section 44. The monitoring section 42and dynamic analysis 44 section outputs are utilized by the gas turbinecontrol system 46 that can send control signals to other gas turbinecontrols subsystems, including industrial gas turbine (IGT) controlssubsystems, such as the fuel injection system 28, in order to unload orshut down the engine 10 in response to changes in monitored combustionconditions within the combustor 20.

Referring to the exemplary controller 40 embodiment shown in FIG. 4, itincludes one or more processors 50, system memory 52 and input/outputcontrol devices 54 for interfacing with the associated engine 10controls, such as the fuel injection control system 28, and the acoustictransceiver/transducer 32, 34 acoustic transmitters and sensors 32 (orfunctionally equivalent performing separate discrete transmitters andreceiver sensors), networks, other computing devices, human machineinterfaces for operator/users, etc. The controller 40 may also includeone or more analog to digital converters 56A and/or other componentsnecessary to allow the controller 40 to interface with the transceivers32, 34 and/or other system components to receive analog sensorinformation. Alternatively, and/or additionally, the system 29 mayinclude one or more analog to digital converters 56B that interfacebetween the transceivers 32, 34 (or functionally equivalent performingseparate discrete transmitters and receiver sensors) and the controller40. As yet a further example, certain transceivers 32, 34 may have ananalog to digital converter 56C integral therewith, or are otherwiseable to communicate digital representations of sensed informationdirectly to the controller 40.

The processor(s) 50 may include one or more processing devices such as ageneral purpose computer, microcomputer or microcontroller. Theprocessors 50 may also comprise one or more processing devices such as acentral processing unit, dedicated digital signal processor (DSP),programmable and/or reprogrammable technology and/or specializedcomponent, such as application specific integrated circuit (ASIC),programmable gate array (e.g., PGA, FPGA).

The memory 52 may include areas for storing computer program codeexecutable by the processor(s) 50, and areas for storing data utilizedfor processing, e.g., memory areas for computing wavelet transforms,Fourier transforms or other executed mathematical operations used tooperate the monitoring and control system 29, as described more fullyherein below. As such, various aspects of the present invention may beimplemented as a computer program product having code configured toperform the detection of combustion engine anomalies of interest,combustion dynamics and engine control functions as set out in greaterdetail herein.

In this regard, the processor(s) 50 and/or memory 52 are programmed withsufficient code, variables, configuration files, etc., to enable thecontroller 40 to perform its designated monitoring and controlfunctions. For example, the controller 40 may be operatively configuredto sense thermoacoustic conditions, analyze thermoacoustic conditionsbased upon inputs from one or more transceiver/transducers 32, 34,control features of the engine 10 in response to its analysis, and/orreport results of its analysis to operators, users, other computerprocesses, etc. as set out in greater detail herein. Thus, all of thedynamic output signals originating from transceiver/transducers 32, 34may be communicated to a single processor 50. In this implementation,the single processor 50 will process the sensor dynamic output signalsusing the data analysis and control functions described in greaterdetail herein, such that it appears as if the results are computed in agenerally parallel fashion. Alternatively, more processors 50 can beused and each processor may be utilized to process one or moretransceiver/transducers 32, 34 dynamic signals, e.g., depending forexample, upon the computation power of each processor.

Monitoring and Control System Operation

The concepts of acoustic temperature and velocity measurements are bothbased on creating a sonic wave, listening to it across the gas streamand finding an average speed of sound across a given path, which is thendescriptive for the gas velocity or velocity/temperature. FIGS. 10 and11 are flow charts illustrating graphically exemplary operation of amonitoring and control system 29 embodiment of the invention thatactively monitors and measures both gas flow velocity and temperatureusing acoustic measurement methodologies. The thick solid and dottedline operational blocks relate to previously described combustiondynamics analysis 42 (solid block), temperature monitoring anddetermination 44 and gas turbine control 46 functions (including by wayof example IGT control functions) that are performed within thecontroller 40. In step 100 sensor signals generated by the sensorcomponents within the transceiver/transducers 32A-H, 34 A-H are read. Instep 110 amplitudes of one or more of the sensor signals are compared topreviously established alarm limits. For example in IGT applications thestep 120 low frequency dynamics (LFD) below 100 Hz are of importancebecause of potential resonance influence at the 50 Hz or 60 Hz enginerotational speed. Other frequency bands of interest are intermediatefrequency dynamics (IFD) between approximately 100-500 Hz and highfrequency dynamics (HFD) above 500 Hz. If an alarm limit is exceeded thecontroller 40 sends a control command, for example to the fuel injectionsystem 28, to unload or shut down the engine 10 in step 400.

If an alarm limit is not exceeded in step 110, then frequency analysisfor dynamics is performed in anomaly detection portion of the combustiondynamics analysis sub system. An exemplary description of how to performanomaly detection is in U.S. Pat. No. 7,853,433 that is incorporatedherein by reference. The sampled high speed dynamic pressure signal isobtained from the sensors in step 130 and time divided into segments instep 140. In step 150 the time-frequency divided sample segments areanalyzed using the wavelet analysis technique described in U.S. Pat. No.7,853,433. Alternatively, a known Fourier spectral analysis thatconverts the time segments into frequency space, analyzes dominantfrequencies by identifying the peak frequencies and their respectiveamplitudes, and identifies amplitudes exceeding defined thresholds. Ifit is determined that a combustion anomaly or anomalies have occurred instep 160 the combustor temperature as determined in the temperaturemonitoring and determination subsystem 44 is compared with the anomalyinformation obtained by the Fourier or wavelet analysis techniques, orboth. In step 180 the anomaly classification as a flame on, flame out orflashback is made in conjunction with the passive or path temperatureinformation obtained from the temperature monitoring and determinationsubsystem 44. For example in a gas turbine flameout the combustortemperature drops off dramatically. Conversely in a flashback scenariothe combustor temperature rises dramatically upstream within thecombustor 14. When the anomaly determination is made in step 180appropriate control signals to unload or shut down the engine are madein the engine control system 46.

The temperature monitoring and determination subsystem 44 may comprisepassive temperature determination utilizing the passive acoustic methoddescribed in United States patent application “Temperature Measurementin a Gas Turbine Engine Combustor, filed on Mar. 14, 2013, Ser. No.13/804,132, incorporated by reference herein, and/or real time actualpath temperature determination within the combustor 14. Real time actualpath temperature is determined by adaptation of the 2-D planar acousticpyrometry technique for gas turbine exhaust system temperaturedetermination described in United States patent Publication No,US2012/0150413 (also incorporated by reference herein) or by a 3-Dtechnique that determines one or more path temperatures between thesensor arrays 32/34 of FIG. 5, that is further described in greaterdetail herein.

In the passive temperature determination method, sampled high speeddynamic pressure signals from the transceiver/transducers 32/34, such asobtained in step 130 are analyzed for dominant modes in step 200.Combustor temperature is calculated based on frequency using the passiveacoustic method in step 210. The passive value is calibrated with areference temperature value in step 220 in order to obtain an activetemperature value within the combustor 14. The calibrated passivetemperature value determined in step 220 is utilized in step 230 todetermine the bulk mean temperature of the combustion gas in step 230.The reference temperature value used in step 220 may be obtained fromone or more thermocouples 36 in the combustor or thermocouples locatedin the exhaust system 18 (not shown). The reference temperature valuemay be an actual path temperature measured in the exhaust system 18, asdescribed in United States Patent Publication No. US2012/0150413 or areal time path temperature measured in the combustor 14 that isdetermined in steps 300-330.

The 2-D real time path temperature is measured by transmitting one ormore acoustic signals in an acoustic transceiver/transducer 32, 34 orother discrete transmitter, such as in the 2-D planar pattern shown forthe (n=8+transceiver/transducers 32A-H in FIG. 9. For example,transceiver/transducer 32A transmits a signal that is received by theremaining (n−1) transceiver/transducers 32B-H and the time-of-flight foreach line-of-sight path is determined. However, at least one, preferablytwo or more sensor elements in the remaining transceiver/transducers32B-H receive the acoustic signal(s) in step 310. Preferably in practiceseveral transceiver/transducers (transmit and receive acoustic signals)circling one plane such that the paths between all transceivers form agrid with desired coarseness which results in the spatial resolution ofthe temperature measurement.

For example, for a cylindrical combustor the transceivers could beequally spaced around the periphery as shown in FIGS. 3 and 9. Thesecould be either fired sequentially (one at a time) or simultaneouslywith disjoint sound patterns that can be readily differentiated. Forsequential firing one transceiver is creating sounds while all remainingtransceivers record it to estimate the travel time for the respectivepaths. Each of these line-of-sight paths represents an averagetemperature along that path. The average temperatures over differentpaths are combined to a two-dimensional map shown in FIG. 9, using aknown computer tomography technique.

The 2-D time-of-flight sound data are converted to gas temperature usingactive acoustics in step 320, such as by utilization of the methodsdescribed in the aforementioned United States Patent Publication No.US2012/0150413 that is incorporated by reference herein. The real timepath temperature that is determined in step 330 is the localized activetemperature value along the line-of-sight transmission path. A pluralityof active temperature values measured along different acoustic paths byperforming the steps 300-330 can be utilized to determine the combustor14 bulk temperatures, alone or in parallel with the dominant frequencypassive acoustic method of steps 200-230. While a single path activetemperature measurement between a single transmitter 30 and acousticsensor 32 provides useful control information, arraying a plurality oftransceiver/transducers 32, 34 selectively in any axial, circumferentialand/or radial pattern or combinations thereof within a combustor 14(see, e.g., FIG. 2, 3, 5, or 9) or in a series of combustors 14facilitates active real time two- or three-dimensional combustiontemperature monitoring within the gas turbine engine 10.

The 2-D or 3-D real time path temperature determined in steps 300-330can be utilized as an input for other monitoring and control functions,with or without one or more of the combustion dynamics analysis 42,passive temperature monitoring and determination 44 and control 46functions described in the exemplary integrated monitoring and controlsystem 29 described herein. For example combustor turbine inlettemperature (TIT) can be actively monitored in real time and used as acontrol parameter for the combustion process. The combustion active pathtemperature determined in steps 300-330 can be utilized to control thefuel/air mixture in the combustor 14 via the fuel injection system 28.The real time path active temperature can be utilized as an input foractive actual gas flow velocity measurement in an industrial gas turbinecombustor or in other types of gas flow environments.

Embodiments of the present invention measure 3-D gas flow velocityand/or gas flow temperature by correlation with sonic time-of-flightalong a line-of-sight sonic pathway between axially spaced, transverselyoriented sonic transmitter and sensor (or transceiver/transducersincorporating the sensors and transmitters), so that the line-of-sightalong the pathway is oriented transverse, as opposed to parallel to thegas flow path. In order to determine gas flow absolute velocity, thetime-of-flight data are corrected or compensated for thermodynamicinfluences on gas temperature, gas constant and speed of sound. As notedabove gas temperature along a line of sight can be determined using thereal time active path temperature or temperature independently obtainedfrom another measurement device (e.g., thermocouple 36). Alternativelylocalized speed of sound c can be determined by measuring bi-directionaltime-of-flight (i.e., forward/downstream transmission andreverse/upstream transmission). The aforementioned thermodynamicinfluences are governed by the known equation:

c(x,y,z)=(γ·R·T)^(1/2)

Where:

c(x,y,z) is the isentropic speed of sound;γ is specific heat ratio;R is the gas constant; andT is the gas temperature.Therefore, once the speed of sound along a path is known, the averagepath temperature and absolute velocity can be determined utilizingembodiments of the invention further described herein.

For accurate absolute velocity or temperature measurement, two planes oftransceiver/transducers 32, 34 are oriented in axially spaced, opposedrelationship within the gas flow, as shown in FIG. 5. The twotransceiver/transducer planes 32, 34 are preferably apart byapproximately the same order of magnitude as the diameter (circular) orwidth (square or rectangular) of the monitored gas flow geometry. Thatis, the axial distance between the two planes should be determinedaccording to the geometry and scale of the interrogated environment aswell as the anticipated or possible ranges of gas flow gas constant,temperature and velocity.

For gas flow velocity estimation, the gas flow is measured axially andtransverse to the flow direction. For example, whentransceiver/transducer 32A in plane Z_(I) fires or transmits a signal,all transceiver/transducers 34B-H in plane Z_(II) that are notparallel-aligned with the signal firing sensor will be listening,thereby creating several paths across the gas flow (n−1 paths for nsensors). The signal transmitting/receiving firing process continuessequentially with the second transceiver/transducer 32B on plane Z_(I)firing to the remaining (n−1) transceiver/transducers 34A and 34C-H,which receive that transmitted signal. The transmitted signal firingwill continue on with the consecutive transceivers firing and creatingn−1 paths for each firing. In the embodiment of FIG. 5, having 8transceivers/transducers in each of the two axially spaced arrays thereare a total of 64 paths in three dimensions. In addition, to alleviatethe directional ambiguity of the velocity (to identify reverse flows andperhaps turbulence fluctuations in reverse direction) the same processwill be repeated with transducer/transceivers 34 in plane Z_(II) firingand transceiver/transducers in plane Z_(I) receiving the reversedirection transmitted acoustic signal, assuming that the gas flowtemperature is already known. Instead of transmitting/firing acousticsignals sequentially from each transceiver/transducer, a sound patternwith a slightly different acoustic signature can be transmitted fromeach respective transceiver/transducer 32A-H, 34A-H simultaneously,which shortens measurement time Referring to steps 500 and 510 of thegas flow velocity measurement method flow chart of FIG. 11, once alltransceiver/transducers in planes Z_(I) and Z_(II) have fired and thetransmitted acoustic signals have been received by the opposing plane oftransversely aligned transceivers/transducers, the process preferablyrepeats continually in real time while a 3-D velocity map u isconstructed from the spatially distributed line-of-sight acoustic paths,using known 3-D tomographic mapping techniques, such as those utilizedin medical or industrial computed tomography systems. The velocityinformation is extracted and mapped, as shown in FIG. 8. Similarly, a3-D temperature map T can be constructed utilizing the time of flightdata, as will be described in greater detail herein.

After all of the transceiver/transducers 32, 34 in a planar array havefired acoustic signals the respective line-of-sight flow pathtime-of-flight data are used to derive absolute velocity in the gas flowpath in step 560, once corrected for the thermodynamic effects oftemperature, gas constant and the speed of sound, as described ingreater detail below. Flow velocity measurement accuracy potentiallydecreases as flow velocity approaches the speed of sound, assumingconstant gas temperature in the velocity measurements. Flow velocitybelow a Mach number of approximately 0.5 is not believed to impactvelocity measurement significantly. Therefore it is preferable, but notrequired, that measured flow velocities should be smaller than half ofthe local speed of sound that is measured. This method can accuratelymeasure high temperature gas flows, including turbine engine gas flows,despite relatively high absolute velocities, because the local speed ofsound increases with temperature.

Once acoustic time-of-flight data are available, they are used by themonitoring and control system 29 or other remote monitoring system todetermine velocity along their respective acoustic paths in accordancewith the remainder of the steps of FIG. 11. Referring to FIGS. 6 and 7,information sound propagation is linearly affected by the gas flow.Relative gas flow velocity for a given temperature, gas constant andspeed of sound is determined by the known equation:

$t_{BC} = {\int_{B}^{C}{\frac{1}{{c( {x,y,z} )} + {{\overset{->}{p}}_{BC} \cdot {\overset{->}{u}( {x,y,z} )}}}{s}}}$

Where:

t_(BC) is the time-of-flight from the first transmitter B to the firstsensor C;

c is the speed of sound in the gas flow for the temperature and gasconstant;

{right arrow over (p)}_(BC) is the unit vector along the first line ofsound path A between B and C; and

{right arrow over (u)}(x,y,z) is velocity vector in the gas flow.

The exemplary planar slice along the line-of-sound path A shows asimplified flow pattern. Referring again to the flow chart of FIG. 11,the relative gas flow velocity is corrected for thermodynamictemperature, gas flow and speed of sound influences, in order to deriveabsolute velocity in step 560. If the path temperature is available(step 520) its influence on the speed of sound can be corrected by knowntomography methods, in order to derive the gas flow absolute velocityalong the line-of-sound path. If the path temperature is not available,times-of-flight for forward (steps 500, 510) and reverse (steps 530,540) acoustic signal transmission are acquired and used to extract thespeed of sound without effect of the gas velocity in accordance with thefollowing equations. The reverse time-of-flight fromtransducer/transceiver C to transducer/transceiver B is determined bythe following equation, similar to that for the forward or downstreamdirection set forth above:

$t_{CB} = {\int_{C}^{B}{\frac{1}{{c( {x,y,z} )} + {{\overset{->}{p}}_{BC} \cdot {\overset{->}{u}( {x,y,z} )}}}{s}}}$

The forward and reverse times-of-flight are added in accordance with thefollowing equation:

${t_{BC} + t_{CB}} = {\int_{B}^{C}{\frac{2 \cdot {c( {x,y,z} )}}{{c( {x,y,z} )}^{2} - {{\overset{->}{p}}_{BC} \cdot {\overset{->}{u}( {x,y,z} )}^{2}}}{s}}}$

Given that the square of the speed of sound c is much greater than thesquare of the gas flow velocity u, the equation is reduced to:

${t_{BC} + t_{CB}} \approx {\int_{B}^{C}{\frac{2}{c( {x,y,z} )}{s}}}$

-   -   where:    -   t_(BC) is the time of flight from the first        transceiver/transducer B to the second transceiver/transducer C;    -   t_(CB) is the time of flight from the second        transceiver/transducer C to the first transceiver/transducer B;    -   c is the speed of sound in the gas flow for the temperature and        gas constant;    -   {right arrow over (p)}_(BC) is the unit vector along the first        line of sound path; and    -   {right arrow over (u)}(x,y,z) is the velocity vector in the gas        flow.

The speed of sound c determined in step 550 of FIG. 11 is then used tocorrect the downstream time-of-flight data for that speed of sound instep 560. The corrected downstream time-of-flight data are used todetermine gas flow absolute velocity in step 570. Where the pathtemperature T along a line-of-flight is not known, the same speed ofsound c determined in step 550 is utilized in some embodiments of theinvention to determine T, using the previously described isentropicspeed of sound relationship

${{c( {x,y,z} )} = ( {\gamma \cdot R \cdot T} )^{\frac{1}{2}}},$

as γ, R and c(x,y,z) are now known. In a similar manner to the pathvelocity determinations previously described, once all the pathtemperatures T are known from each receiver/transmitter unit back andforth, there will be 64 (assuming exemplary 8-sensor case)iso-temperature lines in 3-dimensions. Then using known 3-D tomographicmapping techniques, the 3-dimensional temperature distribution ismapped.

Advantageously the active acoustic temperature and velocity measurementsare performed simultaneously in real time, thus mapping both gas flowtemperature (3-D or alternatively the 2-D mapping of FIG. 9) and 3-D gasflow velocity (FIG. 8). An exemplary acoustic signal transmission andreception timing sequence to perform simultaneous velocity andtemperature measurement is to emit an acoustic signal with atransceiver/transducer on a first array plane (e.g., 32A at Z_(I)). Thecorresponding transversely oriented transceivers/transducers on anaxially spaced opposed second plane (e.g., 34B-H at Z_(II)) receive thesignal for velocity processing and/or temperature processing, if 3-Dtemperature measurement is utilized. If only 2-D temperature measurementis utilized the remainders of the transceiver/transducers on the firstarray plane (e.g., 32B-H at ZI) receive the signal for temperatureprocessing. As previously noted the transmission and receiving processalso can be accelerated by utilizing unique signal transmission patternsfor each transceiver/transducer. There are tradeoffs associated with useof 2-D or 3-D temperature measurement. Where 3-D temperature measurementtechniques are utilized, accuracy of both temperature and velocity mapmay not be the most desired in case of gas velocities of Mach 0.3 orabove as the approximation shown in the equation

${t_{BC} + t_{CB}} \approx {\int_{B}^{C}{\frac{2}{c( {x,y,z} )}{s}}}$

may be less accurate in those velocities ranges, because there are noindependently determined temperature reference values. However,independent temperature T reference values may be determined, using apair of axially separated 2-D acoustic signal sets and two individualacoustic temperature maps determined with the respective 2-Dtime-of-flight signal sets. The 2-D temperature maps are in turninterpolated to create a volumetric temperature map. This volumetric mapwill be used to provide the temperature values T utilized in theisentropic speed of sound equation, along with the known gas constant Rand specific heat ratio γ to extract speeds of sound c. The speed ofsound is then used to extract the velocity vectors u(x,y,z). Once thevelocity vectors are extracted the velocity components can be mapped,eliminating the limitation of below Mach 0.3 gas velocities inherent inthe previously descried 3-D velocity and temperature mapping methods.

Combustor active gas flow velocity or velocity/temperature monitoringutilizing the system and method embodiments described herein with arraysof commonly utilized acoustic sensors is believed to provide fastervelocity and temperature change response than known velocity andtemperature monitoring systems. In accordance with embodiments of theinvention one array of commonly utilized, reliable acoustictransceiver/transducer sensor-transmitters or arrays of separatediscrete acoustic sensors and transmitter pairs can be placed in acombustion flow path under field conditions and monitored to provideactive, real time simultaneous velocity and temperature data and anomalydetection that are all useful for monitoring and control of combustionpower generation equipment, such as industrial gas turbines.

Although various embodiments that incorporate the teachings of thepresent invention have been shown and described in detail herein, thoseskilled in the art can readily devise many other varied embodiments thatstill incorporate these teachings. The invention is not limited in itsapplication to the exemplary embodiment details of construction and thearrangement of components set forth in the description or illustrated inthe drawings. The exemplary engine 10 and exemplary combustor 14 areshown by way of illustration and not by way of limitation, to clearlydescribe certain features and aspects of the present invention set outin greater detail herein. However, the various aspects of the presentinvention described more fully herein may be applied to variouscombustion engines to monitor and/or detect the occurrence of combustionanomalies. The invention is capable of other embodiments and of beingpracticed or of being carried out in various ways. Also, it is to beunderstood that the phraseology and terminology used herein is for thepurpose of description and should not be regarded as limiting. The useof “including,” “comprising,” or “having” and variations thereof hereinis meant to encompass the items listed thereafter and equivalentsthereof as well as additional items. Unless specified or limitedotherwise, the terms “mounted,” “connected,” “supported,” and “coupled”and variations thereof are used broadly and encompass direct andindirect mountings, connections, supports, and couplings. Further,“connected” and “coupled” are not restricted to physical or mechanicalconnections or couplings.

What is claimed is:
 1. A method for actively monitoring gas flowvelocity, comprising: placing at least one first acoustic transmitterand at least one first acoustic sensor respectively oriented in upstreamand downstream transverse locations in a gas flow path and in a distinctfirst line-of-sound path relative to each other, the first sensorcapable of generating first sensor output signals indicative of firstthermoacoustic oscillations; coupling the at least one first transmitterand the at least one first sensor to a controller that is capable ofcausing the at least one first transmitter to transmit first acousticsignals within the gas flow path and capable of correlating first sensoroutput signals with gas flow velocity; transmitting first acousticsignals from the at least one first acoustic transmitter; receiving thefirst acoustic signals from the at least one first acoustic transmitterand generating first dynamic sensor output signals with the at least onefirst acoustic sensor that includes contributions of the received firstacoustic signals; determining a first time-of-flight for the firstacoustic signals traveling along each of the first line of sound paths;and processing the first time-of-flight for the first acoustic signalstraveling along their first line of sound paths to determine respectivegas flow velocity along each respective first line-of-sound path.
 2. Themethod of claim 1, the processing to determine respective absolute gasflow velocity including compensation for gas temperature, gas constantand speed of sound thermodynamic influences on the first time-of-flight.3. The method of claim 2, the temperature compensation comprisingactively monitoring temperatures in the gas flow by: placing at leastone second acoustic transmitter and at least one second acoustic sensorrespectively oriented in a common axial plane in the gas flow path andin a distinct second line-of-sound path relative to each other, thesecond sensor capable of generating second sensor output signalsindicative of second thermoacoustic oscillations; coupling the at leastone second transmitter and the at least one second sensor to thecontroller that is capable of causing the at least one secondtransmitter to transmit second acoustic signals within the gas flow pathand capable of correlating the second sensor output signalstime-of-flight with gas flow temperature; and processing the secondtime-of-flight for the second acoustic signals traveling along thesecond line of sound paths to determine respective gas flow temperaturealong each respective second line-of-sound path.
 4. The method of claim3, the temperature compensation further comprising actively monitoringtemperatures in the gas flow by: placing at least one third acoustictransmitter and at least one third acoustic sensor respectively orientedin a common axial plane in the gas flow path downstream of the secondacoustic transmitter and sensor, and in a distinct third line-of-soundpath relative to each other, the third sensor capable of generatingthird sensor output signals indicative of third thermoacousticoscillations; coupling the at least one third transmitter and the atleast one third sensor to the controller that is capable of causing theat least one third transmitter to transmit third acoustic signals withinthe gas flow path and capable of correlating the third sensor outputsignals time-of-flight with gas flow temperature; processing the thirdtime-of-flight for the third acoustic signals traveling along the thirdline of sound paths to determine respective gas flow temperature alongeach respective third line-of-sound path; and interpolating determinedrespective gas flow temperatures along each respective second and thirdline-of-sound path to create a volumetric temperature map.
 5. The methodof claim 1, the processing of the first time-of-flight described by theequation:$t_{BC} = {\int_{B}^{C}{\frac{1}{{c( {x,y,z} )} + {{\overset{->}{p}}_{BC} \cdot {\overset{->}{u}( {x,y,z} )}}}{s}}}$where: t_(BC) is the time of flight from the first transmitter to thefirst sensor; c is the speed of sound in the gas flow for thetemperature and gas constant; {right arrow over (p)}_(BC) is the unitvector along the first line of sound path; and {right arrow over(u)}(x,y,z) is velocity vector in the gas flow.
 6. The method of claim5, further comprising determining the speed of sound c by: substitutingfor the first transmitters first transceiver/transducers that arecapable of transmitting and receiving acoustic signals, and generatingoutput signals; substituting for the first sensors secondtransceiver/transducers that are capable of transmitting and receivingacoustic signals and generating output signals; coupling the respectiveat least one of the respective first and second transceiver/transducersto the controller that is capable of causing either to transmit firstacoustic signals within the gas flow path and capable of correlatingtransceiver/transducer output signals time-of-flight with gas flowvelocity; transmitting first acoustic signals from the at least onefirst transceiver/transducer; receiving the first acoustic signals fromthe at least one first transceiver/transducer and generating firstdynamic sensor output signals with the at least one secondtransceiver/transducer that includes contributions of the received firstacoustic signals; transmitting reversed first acoustic signals from theat least one second transceiver/transducer; receiving the reversed firstacoustic signals from the at least one second transceiver/transducer andgenerating first reversed dynamic sensor output signals with the atleast one first transceiver/transducer that includes contributions ofthe received reversed first acoustic signals; and determining a firsttime-of-flight for the respective first and first reversed acousticsignals traveling along each of the first line of sound paths; andprocessing the first time-of-flight for the respective first andreversed acoustic signals traveling along their respective first line ofsound paths to determine the speed of sound c.
 7. The method of claim 6,the processing of the respective first time of flight for the respectivefirst and first reversed acoustic signals traveling along each of thefirst line of sound paths to determine the speed of sound c described bythe equation:${t_{BC} + t_{CB}} \approx {\int_{B}^{C}{\frac{2}{c( {x,y,z} )}{s}}}$where: t_(BC) is the time of flight from the firsttransceiver/transducer to the second transceiver/transducer; t_(CB) isthe time of flight from the second transceiver/transducer to the firsttransceiver/transducer; c is the speed of sound in the gas flow for thetemperature and gas constant; {right arrow over (p)}_(BC) is the unitvector along the first line of sound path; and {right arrow over(u)}(x,y,z) is the velocity vector in the gas flow.
 8. The method ofclaim 1, the processing to determine respective absolute gas flowvelocity including compensation for variations in temperature and gasconstant on the speed of sound, comprising: substituting for the firsttransmitters first transceiver/transducers that are capable oftransmitting and receiving acoustic signals, and generating outputsignals; substituting for the first sensors secondtransceiver/transducers that are capable of transmitting and receivingacoustic signals and generating output signals; coupling the respectiveat least one of the respective first and second transceiver/transducersto the controller that is capable of causing either to transmit firstacoustic signals within the gas flow path and capable of correlatingtransceiver/transducer output signals time-of-flight with gas flowvelocity; transmitting first acoustic signals from the at least onefirst transceiver/transducer; receiving the first acoustic signals fromthe at least one first transceiver/transducer and generating firstdynamic sensor output signals with the at least one secondtransceiver/transducer that includes contributions of the received firstacoustic signals; transmitting reversed first acoustic signals from theat least one second transceiver/transducer; receiving the reversed firstacoustic signals from the at least one second transceiver/transducer andgenerating first reversed dynamic sensor output signals with the atleast one first transceiver/transducer that includes contributions ofthe received reversed first acoustic signals; and determining a firsttime-of-flight for the respective first and first reversed acousticsignals traveling along each of the first line of sound paths; andprocessing the first time-of-flight for the respective first andreversed acoustic signals traveling along their respective first line ofsound paths to determine the speed of sound and using the determinedspeed of sound to determine respective absolute gas flow velocity alongeach respective first line-of-sound path.
 9. A method for controllingcombustion in an industrial gas turbine combustor using the gas flowvelocity determined by the method of claim
 1. 10. A system formonitoring gas flow velocity in a turbine engine, comprising: at leastone first acoustic transmitter and at least one first acoustic sensorrespectively oriented in upstream and downstream transverse locations ina gas flow path of a turbine engine and in a distinct firstline-of-sound path relative to each other, the first sensor capable ofgenerating first sensor output signals indicative of firstthermoacoustic oscillations; a controller, coupled to the at least onefirst transmitter and the at least one first sensor that is capable ofcausing the at least one first transmitter to transmit first acousticsignals within the gas flow path and capable of correlating first sensoroutput signals with gas flow velocity, by: transmitting first acousticsignals from the at least one first acoustic transmitter; receiving thefirst acoustic signals from the at least one first acoustic transmitterand generating first dynamic sensor output signals with the at least onefirst acoustic sensor that includes contributions of the received firstacoustic signals; determining a first time-of-flight for the firstacoustic signals traveling along each of the first line of sound paths;and processing the first time-of-flight for the first acoustic signalstraveling along their first line of sound paths to determine respectivegas flow velocity along each respective first line-of-sound path. 11.The system of claim 10, including determination of absolute gas flowvelocity by compensating for temperature thermodynamic influences on thetime-of-flight, further comprising: at least one second acoustictransmitter and at least one second acoustic sensor respectivelyoriented in a common axial plane in the turbine gas flow path and in adistinct second line-of-sound path relative to each other, the secondsensor capable of generating second sensor output signals indicative ofsecond thermoacoustic oscillations; the controller coupled to the atleast one second transmitter and the at least one second sensor andcapable of: causing the at least one second transmitter to transmitsecond acoustic signals within the gas flow path; and processing thesecond time-of-flight for the second acoustic signals traveling alongthe second line of sound paths to determine respective gas flowtemperature along each respective second line-of-sound path; and usingthe determined respective gas flow temperatures in compensating fortemperature influence on the first time-of-flight.
 12. The system ofclaim 11 further comprising first and second planar arrays of ntransceiver/transducers incorporating the respective first and secondtransmitters and sensors, the planar arrays oriented generally normal tomutually spaced along the gas flow path in opposed relationship, eachtransceiver/transducer in the first planar array transmitting: the firstacoustic signals to (n−1) transceiver/transducers in the second planararray; and the second acoustic signals to (n−1) transceiver/transducersin the first planar array.
 13. The system of claim 10, the controllerprocessing of the first time-of-flight described by the equation:$t_{BC} = {\int_{B}^{C}{\frac{1}{{c( {x,y,z} )} + {{\overset{->}{p}}_{BC} \cdot {\overset{->}{u}( {x,y,z} )}}}{s}}}$where: t_(BC) is the time-of-flight from the first transmitter to thefirst sensor; c is the speed of sound in the gas flow for thetemperature and gas constant; {right arrow over (p)}_(BC) is the unitvector along the first line of sound path; and {right arrow over(u)}(x,y,z) is velocity vector in the gas flow.
 14. The system of claim13, further comprising the controller determining the speed of sound cby: substituting for the first transmitters firsttransceiver/transducers that are capable of transmitting and receivingacoustic signals, and generating output signals; substituting for thefirst sensors second transceiver/transducers that are capable oftransmitting and receiving acoustic signals and generating outputsignals; coupling the respective at least one of the respective firstand second transceiver/transducers to the controller that is capable ofcausing either to transmit first acoustic signals within the gas flowpath and capable of correlating transceiver/transducer outputtime-of-flight signals with gas flow velocity; transmitting firstacoustic signals from the at least one first transceiver/transducer;receiving the first acoustic signals from the at least one firsttransceiver/transducer and generating first dynamic sensor outputsignals with the at least one second transceiver/transducer thatincludes contributions of the received first acoustic signals;transmitting reversed first acoustic signals from the at least onesecond transceiver/transducer; receiving the reversed first acousticsignals from the at least one second transceiver/transducer andgenerating first reversed dynamic sensor output signals with the atleast one first transceiver/transducer that includes contributions ofthe received reversed first acoustic signals; and determining a firsttime-of-flight for the respective first and first reversed acousticsignals traveling along each of the first line of sound paths; andprocessing the first time-of-flight for the respective first andreversed acoustic signals traveling along their respective first line ofsound paths to determine the speed of sound c.
 15. The system of claim14, the controller processing of the respective first time of flight forthe respective first and first reversed acoustic signals traveling alongeach of the first line of sound paths to determine the speed of sound cdescribed by the equation:${t_{BC} + t_{CB}} \approx {\int_{B}^{C}{\frac{2}{c( {x,y,z} )}{s}}}$where: t_(BC) is the time-of-flight from the firsttransceiver/transducer to the second transceiver/transducer; t_(CB) isthe time-of-flight from the second transceiver/transducer to the firsttransceiver/transducer; c is the speed of sound in the gas flow for thetemperature and gas constant; {right arrow over (p)}_(BC) is the unitvector along the first line of sound path; and {right arrow over(u)}(x,y,z) is the velocity vector in the gas flow.
 16. A gas turbineengine apparatus, comprising: a compressor section; a combustor sectionincluding a plurality of combustors, each combustor having an injectorsystem for regulating fuel/air mixture; a turbine section; and a systemfor monitoring gas flow velocity in the turbine engine, comprising: atleast one first acoustic transmitter and at least one first acousticsensor respectively oriented in upstream and downstream transverselocations in a gas flow path of a turbine engine and in a distinct firstline-of-sound path relative to each other, the first sensor capable ofgenerating first sensor output signals indicative of firstthermoacoustic oscillations; a controller, coupled to the at least onefirst transmitter and the at least one first sensor that is capable ofcausing the at least one first transmitter to transmit first acousticsignals within the gas flow path and capable of correlating first sensoroutput signals with gas flow velocity, by: transmitting first acousticsignals from the at least one first acoustic transmitter; receiving thefirst acoustic signals from the at least one first acoustic transmitterand generating first dynamic sensor output signals with the at least onefirst thermoacoustic sensor that includes contributions of the receivedfirst acoustic signals; determining a first time-of-flight for the firstacoustic signals traveling along each of the first line of sound paths;and processing the first time-of-flight for the first acoustic signalstraveling along their first line of sound paths to determine respectivegas flow velocity along each respective first line-of-sound path. 17.The apparatus of claim 16, including determination of absolute gas flowvelocity by compensating for temperature thermodynamic influences on thetime-of-flight, further comprising: at least one second acoustictransmitter and at least one second acoustic sensor respectivelyoriented in a common axial plane in the turbine gas flow path and in adistinct second line-of-sound path relative to each other, the secondsensor capable of generating second sensor output signals indicative ofsecond thermoacoustic oscillations; the controller coupled to the atleast one second transmitter and the at least one second sensor andcapable of: causing the at least one second transmitter to transmitsecond acoustic signals within the gas flow path; and processing thesecond time-of-flight for the second acoustic signals traveling alongthe second line of sound paths to determine respective gas flowtemperature along each respective second line-of-sound path; and usingthe determined respective gas flow temperatures in compensating fortemperature on the first time-of-flight.
 18. The apparatus of claim 17,further comprising first and second planar arrays of ntransceiver/transducers incorporating the respective first and secondtransmitters and sensors, the planar arrays oriented generally normal toand mutually spaced along the gas flow path in opposed relationship,each transceiver/transducer in the first planar array transmitting: thefirst acoustic signals to (n−1) transceiver/transducers in the secondplanar array; and the second acoustic signals to (n−1)transceiver/transducers in the first planar array.
 19. The system ofclaim 16, the controller processing of the first time-of-flightdescribed by the equation:$t_{BC} = {\int_{B}^{C}{\frac{1}{{c( {x,y,z} )} + {{\overset{->}{p}}_{BC} \cdot {\overset{->}{u}( {x,y,z} )}}}{s}}}$where: t_(BC) is the time of flight from the first transmitter to thefirst sensor; c is the speed of sound in the gas flow for thetemperature and gas constant; {right arrow over (p)}_(BC) is the unitvector along the first line of sound path; and {right arrow over(u)}(x,y,z) is velocity vector in the gas flow.
 20. The apparatus ofclaim 19, further comprising the controller determining the speed ofsound c by: substituting for the first transmitters firsttransceiver/transducers that are capable of transmitting and receivingacoustic signals, and generating output signals; substituting for thefirst sensors second transceiver/transducers that are capable oftransmitting and receiving acoustic signals and generating outputsignals; coupling the respective at least one of the respective firstand second transceiver/transducers to the controller that is capable ofcausing either to transmit first acoustic signals within the gas flowpath and capable of correlating transceiver/transducer output signalstime-of-flight with gas flow velocity; transmitting first acousticsignals from the at least one first transceiver/transducer; receivingthe first acoustic signals from the at least one firsttransceiver/transducer and generating first dynamic sensor outputsignals with the at least one second transceiver/transducer thatincludes contributions of the received first acoustic signals;transmitting reversed first acoustic signals from the at least onesecond transceiver/transducer; receiving the reversed first acousticsignals from the at least one second transceiver/transducer andgenerating first reversed dynamic sensor output signals with the atleast one first transceiver/transducer that includes contributions ofthe received reversed first acoustic signals; and determining a firsttime-of-flight for the respective first and first reversed acousticsignals traveling along each of the first line of sound paths; andprocessing the first time-of-flight for the respective first andreversed acoustic signals traveling along their respective first line ofsound paths to determine the speed of sound c, described by theequation:${t_{BC} + t_{CB}} \approx {\int_{B}^{C}{\frac{2}{c( {x,y,z} )}{s}}}$where: t_(BC) is the time of flight from the firsttransceiver/transducer to the second transceiver/transducer; t_(CB) isthe time of flight from the second transceiver/transducer to the firsttransceiver/transducer; c is the speed of sound in the gas flow for thetemperature and gas constant; {right arrow over (p)}_(BC) is the unitvector along the first line of sound path; and {right arrow over(u)}(x,y,z) is the velocity vector in the gas flow.